US20160032738A1 - Gas turbine engine blade tip treatment - Google Patents
Gas turbine engine blade tip treatment Download PDFInfo
- Publication number
- US20160032738A1 US20160032738A1 US14/803,497 US201514803497A US2016032738A1 US 20160032738 A1 US20160032738 A1 US 20160032738A1 US 201514803497 A US201514803497 A US 201514803497A US 2016032738 A1 US2016032738 A1 US 2016032738A1
- Authority
- US
- United States
- Prior art keywords
- fan blade
- ceramic
- oxidation layer
- blade body
- tip
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/288—Protective coatings for blades
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B05—SPRAYING OR ATOMISING IN GENERAL; APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
- B05D—PROCESSES FOR APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
- B05D1/00—Processes for applying liquids or other fluent materials
- B05D1/18—Processes for applying liquids or other fluent materials performed by dipping
-
- C—CHEMISTRY; METALLURGY
- C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
- C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
- C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
-
- C—CHEMISTRY; METALLURGY
- C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
- C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
- C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
- C23C28/04—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D only coatings of inorganic non-metallic material
-
- C—CHEMISTRY; METALLURGY
- C25—ELECTROLYTIC OR ELECTROPHORETIC PROCESSES; APPARATUS THEREFOR
- C25D—PROCESSES FOR THE ELECTROLYTIC OR ELECTROPHORETIC PRODUCTION OF COATINGS; ELECTROFORMING; APPARATUS THEREFOR
- C25D11/00—Electrolytic coating by surface reaction, i.e. forming conversion layers
- C25D11/02—Anodisation
- C25D11/04—Anodisation of aluminium or alloys based thereon
-
- C—CHEMISTRY; METALLURGY
- C25—ELECTROLYTIC OR ELECTROPHORETIC PROCESSES; APPARATUS THEREFOR
- C25D—PROCESSES FOR THE ELECTROLYTIC OR ELECTROPHORETIC PRODUCTION OF COATINGS; ELECTROFORMING; APPARATUS THEREFOR
- C25D11/00—Electrolytic coating by surface reaction, i.e. forming conversion layers
- C25D11/02—Anodisation
- C25D11/26—Anodisation of refractory metals or alloys based thereon
-
- C—CHEMISTRY; METALLURGY
- C25—ELECTROLYTIC OR ELECTROPHORETIC PROCESSES; APPARATUS THEREFOR
- C25D—PROCESSES FOR THE ELECTROLYTIC OR ELECTROPHORETIC PRODUCTION OF COATINGS; ELECTROFORMING; APPARATUS THEREFOR
- C25D13/00—Electrophoretic coating characterised by the process
- C25D13/02—Electrophoretic coating characterised by the process with inorganic material
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
- F04D29/324—Blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/90—Coating; Surface treatment
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/20—Oxide or non-oxide ceramics
- F05D2300/21—Oxide ceramics
- F05D2300/2112—Aluminium oxides
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/605—Crystalline
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/611—Coating
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10S977/00—Nanotechnology
- Y10S977/70—Nanostructure
- Y10S977/773—Nanoparticle, i.e. structure having three dimensions of 100 nm or less
- Y10S977/775—Nanosized powder or flake, e.g. nanosized catalyst
- Y10S977/776—Ceramic powder or flake
Definitions
- the present disclosure relates generally to a gas turbine engine, and more specifically to a blade tip treatment for a gas turbine engine.
- fan blade includes an aluminum substrate or fan blade body.
- a polyurethane coating is applied over the fan blade body.
- an outer air seal is provided in the form of an abradable rub strip.
- a tip of the fan blade may rub against the adjacent outer air seal rub strip.
- the tip may wear and heat may be generated.
- the tip of the fan blade body has been anodized to create a hard coating layer.
- Heat is generated when the blade hard coating rubs against the abradable rub strip. Because of the very low thermal conductivity ( ⁇ 0.1 W/m K) of a typical abradable rub strip material as compared to the hard coating layer conductivity of ⁇ 30 W/m K and the aluminum fan body conductivity of ⁇ 160 W/m K, most of the heat generated during a rub event is conducted into the fan blade. Such conduction can cause the blade temperature near the tip to exceed the capability of the polyurethane, and results in degradation of the bonding of polyurethane to the aluminum blade.
- a method of manufacturing a fan blade comprising: immersing at least a crystalline oxidation layer of a metallic fan blade body in a solution of ceramic nanosheets in suspension, the ceramic nanosheets having a charge of a first polarity; and applying a potential of a second polarity to the fan blade body while the crystalline oxidation layer is immersed in the solution of ceramic nanosheets in suspension, wherein the second polarity is different than the first polarity.
- the fan blade body includes a tip, the tip including the crystalline oxidation layer.
- the crystalline oxidation layer includes a plurality of pore channels therein; and said applying a potential electrophoretically drives at least some of the ceramic nanosheets into at least some of the pore channels.
- the crystalline oxidation layer is an aluminum oxide hard coating layer.
- the aluminum oxide hard coating layer is a MIL-A-8625F Type III coating.
- the ceramic nanosheets comprise a ceramic selected from the group consisting of: MoS 2 , MoSe 2 , WS 2 , MoSi 2 , WSe 2 , TiS 2 , TaS 2 , and ZrS 2 .
- the first polarity is negative and the second polarity is positive.
- the potential is a DC potential.
- the fan blade body comprises one of a 7000 series and a 2000 series aluminum alloy.
- the ceramic nanosheets are formed by a process comprising the steps of: intercalating lithium cations between layers of a ceramic starting material to produce an intercalated ceramic structure; and exfoliating the intercalated ceramic structure to produce the ceramic nanosheets.
- the adhering step includes arranging an adhesive-saturated scrim between the sheath and the leading edge.
- the crystalline oxidation layer is left exposed subsequent to the coating step.
- a fan blade for a gas turbine engine comprising: a metallic fan blade body including a tip with a crystalline oxidation layer; wherein the crystalline oxidation layer includes a plurality of pore channels; and wherein at least some of the plurality of pore channels include ceramic nanosheets.
- the fan blade body includes a tip, the tip including the crystalline oxidation layer.
- the crystalline oxidation layer is an aluminum oxide hard coating layer.
- the aluminum oxide hard coating layer is a MIL-A-8625F Type III coating.
- the ceramic nanosheets comprise a ceramic selected from the group consisting of: MoS 2 , MoSe 2 , WS 2 , MoSi 2 , WSe 2 , TiS 2 , TaS 2 , and ZrS 2 .
- the metallic fan blade body comprises one of a 7000 series and a 2000 series aluminum alloy.
- FIG. 1 is a schematic partial cross-sectional view of a gas turbine engine in an embodiment.
- FIG. 2 is a perspective view of an embodiment of a fan blade of the engine shown in FIG. 1 .
- FIG. 3 is an end view of the fan blade shown in FIG. 2 , according to an embodiment.
- FIG. 4 is a schematic cross-sectional view of a fan blade tip and hard coating layer in an embodiment.
- FIG. 5 is a schematic process diagram of a lithium intercalation and exfoliation process in an embodiment.
- FIG. 6 is a schematic cross-sectional view of a fan blade tip and hard coating layer in an embodiment.
- FIG. 7 is a schematic cross-sectional view of a fan blade tip and hard coating layer in an embodiment.
- FIG. 1 schematically illustrates a gas turbine engine 20 .
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- Alternative engines might include an augmentor section (not shown) among other systems or features.
- the fan section 22 drives air along a bypass flow path B in a bypass duct, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
- the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 . It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
- the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42 , a low pressure compressor 44 and a low pressure turbine 46 .
- the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30 .
- the high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54 .
- a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54 .
- An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
- the engine static structure 36 further supports bearing systems 38 in the turbine section 28 .
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- each of the positions of the fan section 22 , compressor section 24 , combustor section 26 , turbine section 28 , and fan drive gear system 48 may be varied.
- gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28
- fan section 22 may be positioned forward or aft of the location of gear system 48 .
- the engine 20 in one example is a high-bypass geared aircraft engine.
- the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10)
- the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3
- the low pressure turbine 46 has a pressure ratio that is greater than about five.
- the engine 20 bypass ratio is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1.
- Low pressure turbine 46 pressure ratio is the pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
- the fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters).
- TSFC Thrust Specific Fuel Consumption
- Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
- Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 .
- the “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).
- a fan blade 60 of the fan 42 includes a root 62 supporting a platform 64 .
- An airfoil 66 extends from the platform 64 to a tip 67 .
- the airfoil 66 includes spaced apart leading and trailing edges 68 , 70 .
- Pressure and suction sides 72 ( FIG. 3 ), 74 adjoin the leading and trailing edges 68 , 70 to provide a fan blade contour 86 .
- the fan blade 60 is exemplary, and “platformless” fan blades may also be used.
- the tip 67 is arranged adjacent to a sealing structure 83 , which is typically arranged in relation to the tip 67 to provide a clearance 84 .
- the tip 67 may be prone to rubbing the sealing structure 83 , which can generate heat and undesirably wear the tip 67 .
- Each fan blade 60 includes an aluminum fan blade body 80 , which may be hollow or solid.
- the fan blade body 80 is constructed from a 7000 series aluminum alloy, such as 7255. In another example, a 2000 series aluminum is used. Other aluminum alloys, other metals, and other materials such as, for example, titanium and titanium alloys, may also be used.
- the fan blade body 80 has a leading edge 78 .
- a sheath 76 may be secured to the fan blade body 80 over the edge 78 with adhesive 82 .
- the sheath 76 and the fan blade body 80 are constructed from first and second metals that are different from one another.
- the sheath 76 is constructed from a titanium alloy. It should be understood that other metals or materials may be used.
- the adhesive 82 provides a barrier between the fan blade body 80 and the sheath 76 to prevent galvanic corrosion. Referring to FIG. 3 , the adhesive 82 may include a scrim 88 (e. g., a glass fiber scrim) that carries the adhesive 82 in an embodiment.
- a polymer coating 90 may be applied over the fan blade body 80 adjacent to the sheath 76 to provide a fan blade contour 86 .
- the coating 90 is polyurethane in one embodiment, but may be any other polymeric coating.
- the substrate provided by the fan blade body 80 may be masked to leave at least the tip 67 exposed so that the tip 67 may be oxidized.
- a chemical masking, mechanical masking tape, lacquer, or other painted on coating may be used and temporarily deposited upon the exterior surface of the fan blade body 80 . Since the oxidation process requires significant voltage, reducing the area to be oxidized greatly reduces the necessary power and cost of oxidizing the tip 67 .
- the tip 67 is oxidized to produce a crystalline aluminum oxide hard coating layer 92 , as shown in FIG. 3 .
- the hard coating layer 92 may be a Type III hard anodic coating as described in United States Military Specification MIL-A-8625F.
- the hard coating layer 92 has a porous structure with pore channels 94 aligned substantially perpendicular to the hard coating layer 92 surface 96 .
- ceramic nanosheets may be deposited inside the pore channels 94 of the hard coating layer 92 using, for example, electrophoretic force. The ceramic nanosheets will reduce the thermal conductivity of the hard coating layer 92 without changing the mechanical properties of the hard coating layer 92 since the ceramic nanosheets do not comprise a layer applied to the surface 96 of the hard coating layer 92 .
- ceramic nanosheets may be generated through an intercalation and exfoliation process, to name just one non-limiting example, which generates a ceramic nanosheet suspension stabilized with a negative electrical charge on the surface.
- Starting ceramic materials 100 such as MoS 2 , MoSe 2 , WS 2 , MoSi 2 , WSe 2 , TiS 2 , TaS 2 , and ZrS 2 , to name just a few non-limiting examples, may be used.
- a process 102 of lithium intercalation may be used to intercalate lithium cations between the layers of starting ceramic material 100 in a liquid environment, swelling the crystal and weakening the interlayer attraction within the starting ceramic material 100 .
- lithium intercalation processes including the N-butyl lithium method, the lithium metal dissolved in liquid ammonia method, and the electrochemical lithiation method, which provides high yield single layer ceramic nanosheets, to name just a few non-limiting examples.
- the present embodiments include the use of any intercalation process. This produces the intercalated ceramic structure 104 .
- the lithium cations may then be removed via an exfoliation process 106 in water, where lithium cations are surrounded by H 2 O molecules (shown at 106 ) and exfoliated from the intercalated ceramic structure 104 , producing a suspension of ceramic nanosheets 108 having negative electrical charge 110 on the surface thereof.
- the hard coating layer 92 may then be submerged in the ceramic nanosheet 108 suspension.
- a positive potential 112 such as a positive DC potential in an embodiment, may be connected to the fan blade body 80 to electrophoretically drive the ceramic nanosheets 108 into the pore channels 94 of the hard coating layer 92 , as shown in FIGS. 6 and 7 . It will be appreciated from the present disclosure that if the ceramic nanosheets 108 have a positive charge, then a negative charge may be applied to the fan blade body 80 .
- the ceramic nanosheets 108 have an extremely low coefficient of friction and will also serve as a lubricant during rubbing between the tip 67 and the sealing structure 83 . With less friction during rub events, less heat will be generated during abrasion.
Abstract
Description
- This application claims the benefit of and incorporates by reference herein the disclosure of U.S. Ser. No. 62/032,856, filed Aug. 4, 2015.
- The present disclosure relates generally to a gas turbine engine, and more specifically to a blade tip treatment for a gas turbine engine.
- Many turbine engines have fans formed by a plurality of blades. One example fan blade includes an aluminum substrate or fan blade body. A polyurethane coating is applied over the fan blade body. In order to obtain better fan performance, an outer air seal is provided in the form of an abradable rub strip. During operation of the gas turbine engine, a tip of the fan blade may rub against the adjacent outer air seal rub strip. During the rub event, the tip may wear and heat may be generated. To this end, the tip of the fan blade body has been anodized to create a hard coating layer.
- Heat is generated when the blade hard coating rubs against the abradable rub strip. Because of the very low thermal conductivity (˜0.1 W/m K) of a typical abradable rub strip material as compared to the hard coating layer conductivity of ˜30 W/m K and the aluminum fan body conductivity of ˜160 W/m K, most of the heat generated during a rub event is conducted into the fan blade. Such conduction can cause the blade temperature near the tip to exceed the capability of the polyurethane, and results in degradation of the bonding of polyurethane to the aluminum blade.
- In one embodiment, a method of manufacturing a fan blade is disclosed comprising: immersing at least a crystalline oxidation layer of a metallic fan blade body in a solution of ceramic nanosheets in suspension, the ceramic nanosheets having a charge of a first polarity; and applying a potential of a second polarity to the fan blade body while the crystalline oxidation layer is immersed in the solution of ceramic nanosheets in suspension, wherein the second polarity is different than the first polarity.
- In a further embodiment of the above, the fan blade body includes a tip, the tip including the crystalline oxidation layer.
- In a further embodiment of any of the above, the crystalline oxidation layer includes a plurality of pore channels therein; and said applying a potential electrophoretically drives at least some of the ceramic nanosheets into at least some of the pore channels.
- In a further embodiment of any of the above, the crystalline oxidation layer is an aluminum oxide hard coating layer.
- In a further embodiment of any of the above, the aluminum oxide hard coating layer is a MIL-A-8625F Type III coating.
- In a further embodiment of any of the above, the ceramic nanosheets comprise a ceramic selected from the group consisting of: MoS2, MoSe2, WS2, MoSi2, WSe2, TiS2, TaS2, and ZrS2.
- In a further embodiment of any of the above, the first polarity is negative and the second polarity is positive.
- In a further embodiment of any of the above, the potential is a DC potential.
- In a further embodiment of any of the above, the fan blade body comprises one of a 7000 series and a 2000 series aluminum alloy.
- In a further embodiment of any of the above, the ceramic nanosheets are formed by a process comprising the steps of: intercalating lithium cations between layers of a ceramic starting material to produce an intercalated ceramic structure; and exfoliating the intercalated ceramic structure to produce the ceramic nanosheets.
- In a further embodiment of any of the above, further comprising the step of adhering a sheath to a leading edge of the fan blade body.
- In a further embodiment of any of the above, the adhering step includes arranging an adhesive-saturated scrim between the sheath and the leading edge.
- In a further embodiment of any of the above, further comprising the step of coating the fan blade body with polyurethane to provide a fan blade contour along with the sheath.
- In a further embodiment of any of the above, the crystalline oxidation layer is left exposed subsequent to the coating step.
- In another embodiment, a fan blade for a gas turbine engine is disclosed comprising: a metallic fan blade body including a tip with a crystalline oxidation layer; wherein the crystalline oxidation layer includes a plurality of pore channels; and wherein at least some of the plurality of pore channels include ceramic nanosheets.
- In a further embodiment of the above, the fan blade body includes a tip, the tip including the crystalline oxidation layer.
- In a further embodiment of any of the above, the crystalline oxidation layer is an aluminum oxide hard coating layer.
- In a further embodiment of any of the above, the aluminum oxide hard coating layer is a MIL-A-8625F Type III coating.
- In a further embodiment of any of the above, the ceramic nanosheets comprise a ceramic selected from the group consisting of: MoS2, MoSe2, WS2, MoSi2, WSe2, TiS2, TaS2, and ZrS2.
- In a further embodiment of any of the above, the metallic fan blade body comprises one of a 7000 series and a 2000 series aluminum alloy.
- Other embodiments are also disclosed.
- The embodiments and other features, advantages and disclosures contained herein, and the manner of attaining them, will become apparent and the present disclosure will be better understood by reference to the following description of various exemplary embodiments of the present disclosure taken in conjunction with the accompanying drawings, wherein:
-
FIG. 1 is a schematic partial cross-sectional view of a gas turbine engine in an embodiment. -
FIG. 2 is a perspective view of an embodiment of a fan blade of the engine shown inFIG. 1 . -
FIG. 3 is an end view of the fan blade shown inFIG. 2 , according to an embodiment. -
FIG. 4 is a schematic cross-sectional view of a fan blade tip and hard coating layer in an embodiment. -
FIG. 5 is a schematic process diagram of a lithium intercalation and exfoliation process in an embodiment. -
FIG. 6 is a schematic cross-sectional view of a fan blade tip and hard coating layer in an embodiment. -
FIG. 7 is a schematic cross-sectional view of a fan blade tip and hard coating layer in an embodiment. - For the purposes of promoting an understanding of the principles of the invention, reference will now be made to certain embodiments and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended, and alterations and modifications in the illustrated device, and further applications of the principles of the invention as illustrated therein are herein contemplated as would normally occur to one skilled in the art to which the invention relates.
-
FIG. 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B in a bypass duct, while thecompressor section 24 drives air along a core flow path C for compression and communication into thecombustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. - The
exemplary engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing systems 38. It should be understood thatvarious bearing systems 38 at various locations may alternatively or additionally be provided, and the location ofbearing systems 38 may be varied as appropriate to the application. - The
low speed spool 30 generally includes aninner shaft 40 that interconnects afan 42, alow pressure compressor 44 and alow pressure turbine 46. Theinner shaft 40 is connected to thefan 42 through a speed change mechanism, which in exemplarygas turbine engine 20 is illustrated as a gearedarchitecture 48 to drive thefan 42 at a lower speed than thelow speed spool 30. Thehigh speed spool 32 includes anouter shaft 50 that interconnects ahigh pressure compressor 52 andhigh pressure turbine 54. Acombustor 56 is arranged inexemplary gas turbine 20 between thehigh pressure compressor 52 and thehigh pressure turbine 54. An enginestatic structure 36 is arranged generally between thehigh pressure turbine 54 and thelow pressure turbine 46. The enginestatic structure 36 further supports bearingsystems 38 in theturbine section 28. Theinner shaft 40 and theouter shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes. - The core airflow is compressed by the
low pressure compressor 44 then thehigh pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over thehigh pressure turbine 54 andlow pressure turbine 46. Theturbines low speed spool 30 andhigh speed spool 32 in response to the expansion. It will be appreciated that each of the positions of thefan section 22,compressor section 24,combustor section 26,turbine section 28, and fandrive gear system 48 may be varied. For example,gear system 48 may be located aft ofcombustor section 26 or even aft ofturbine section 28, andfan section 22 may be positioned forward or aft of the location ofgear system 48. - The
engine 20 in one example is a high-bypass geared aircraft engine. In a further example, theengine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and thelow pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, theengine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 44, and thelow pressure turbine 46 has a pressure ratio that is greater than about five 5:1.Low pressure turbine 46 pressure ratio is the pressure measured prior to inlet oflow pressure turbine 46 as related to the pressure at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle. The gearedarchitecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. - A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 22 of theengine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec). - Referring to
FIG. 2 , afan blade 60 of thefan 42 includes aroot 62 supporting aplatform 64. Anairfoil 66 extends from theplatform 64 to atip 67. Theairfoil 66 includes spaced apart leading and trailingedges FIG. 3 ), 74 adjoin the leading and trailingedges fan blade contour 86. It should be understood that thefan blade 60 is exemplary, and “platformless” fan blades may also be used. - The
tip 67 is arranged adjacent to a sealingstructure 83, which is typically arranged in relation to thetip 67 to provide aclearance 84. During certain engine operating conditions, thetip 67 may be prone to rubbing the sealingstructure 83, which can generate heat and undesirably wear thetip 67. - Each
fan blade 60 includes an aluminumfan blade body 80, which may be hollow or solid. In one example, thefan blade body 80 is constructed from a 7000 series aluminum alloy, such as 7255. In another example, a 2000 series aluminum is used. Other aluminum alloys, other metals, and other materials such as, for example, titanium and titanium alloys, may also be used. - The
fan blade body 80 has aleading edge 78. Asheath 76 may be secured to thefan blade body 80 over theedge 78 withadhesive 82. In one example, thesheath 76 and thefan blade body 80 are constructed from first and second metals that are different from one another. In one example, thesheath 76 is constructed from a titanium alloy. It should be understood that other metals or materials may be used. The adhesive 82 provides a barrier between thefan blade body 80 and thesheath 76 to prevent galvanic corrosion. Referring toFIG. 3 , the adhesive 82 may include a scrim 88 (e. g., a glass fiber scrim) that carries the adhesive 82 in an embodiment. - A
polymer coating 90 may be applied over thefan blade body 80 adjacent to thesheath 76 to provide afan blade contour 86. Thecoating 90 is polyurethane in one embodiment, but may be any other polymeric coating. The substrate provided by thefan blade body 80 may be masked to leave at least thetip 67 exposed so that thetip 67 may be oxidized. A chemical masking, mechanical masking tape, lacquer, or other painted on coating may be used and temporarily deposited upon the exterior surface of thefan blade body 80. Since the oxidation process requires significant voltage, reducing the area to be oxidized greatly reduces the necessary power and cost of oxidizing thetip 67. - The
tip 67 is oxidized to produce a crystalline aluminum oxidehard coating layer 92, as shown inFIG. 3 . For example, thehard coating layer 92 may be a Type III hard anodic coating as described in United States Military Specification MIL-A-8625F. - Referring to the schematic representation of
FIG. 4 , in an embodiment, thehard coating layer 92 has a porous structure withpore channels 94 aligned substantially perpendicular to thehard coating layer 92surface 96. In order to reduce the thermal conductivity of thehard coating layer 92, ceramic nanosheets may be deposited inside thepore channels 94 of thehard coating layer 92 using, for example, electrophoretic force. The ceramic nanosheets will reduce the thermal conductivity of thehard coating layer 92 without changing the mechanical properties of thehard coating layer 92 since the ceramic nanosheets do not comprise a layer applied to thesurface 96 of thehard coating layer 92. - Referring to
FIG. 5 , ceramic nanosheets may be generated through an intercalation and exfoliation process, to name just one non-limiting example, which generates a ceramic nanosheet suspension stabilized with a negative electrical charge on the surface. Startingceramic materials 100, such as MoS2, MoSe2, WS2, MoSi2, WSe2, TiS2, TaS2, and ZrS2, to name just a few non-limiting examples, may be used. Aprocess 102 of lithium intercalation may be used to intercalate lithium cations between the layers of startingceramic material 100 in a liquid environment, swelling the crystal and weakening the interlayer attraction within the startingceramic material 100. There are multiple lithium intercalation processes known in the art, including the N-butyl lithium method, the lithium metal dissolved in liquid ammonia method, and the electrochemical lithiation method, which provides high yield single layer ceramic nanosheets, to name just a few non-limiting examples. The present embodiments include the use of any intercalation process. This produces the intercalatedceramic structure 104. The lithium cations may then be removed via anexfoliation process 106 in water, where lithium cations are surrounded by H2O molecules (shown at 106) and exfoliated from the intercalatedceramic structure 104, producing a suspension ofceramic nanosheets 108 having negativeelectrical charge 110 on the surface thereof. - The
hard coating layer 92 may then be submerged in theceramic nanosheet 108 suspension. Apositive potential 112, such as a positive DC potential in an embodiment, may be connected to thefan blade body 80 to electrophoretically drive theceramic nanosheets 108 into thepore channels 94 of thehard coating layer 92, as shown inFIGS. 6 and 7 . It will be appreciated from the present disclosure that if theceramic nanosheets 108 have a positive charge, then a negative charge may be applied to thefan blade body 80. Theceramic nanosheets 108 have an extremely low coefficient of friction and will also serve as a lubricant during rubbing between thetip 67 and the sealingstructure 83. With less friction during rub events, less heat will be generated during abrasion. With theceramic nanosheets 108 filling the hard coating, phonon scattering will increase and thus the thermal conductivity of thehard coating layer 92 will decrease. The addition of theceramic nanosheets 108 to thehard coating layer 92 does not change the mechanical properties of thehard coating layer 92 with respect to abrasion resistance, and the added lubrication provided by theceramic nanosheets 108 will reduce the wear of thehard coating layer 92. Both the lower friction and reduced thermal conductivity of thehard coating layer 92 operate to lower thefan blade body 80 temperature rise during rub events and reduce the risk of compromising thepolyurethane coating 90 on thefan blade body 80. - While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only certain embodiments have been shown and described and that all changes and modifications that come within the spirit of the invention are desired to be protected.
Claims (20)
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US14/803,497 US10094227B2 (en) | 2014-08-04 | 2015-07-20 | Gas turbine engine blade tip treatment |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US201462032856P | 2014-08-04 | 2014-08-04 | |
US14/803,497 US10094227B2 (en) | 2014-08-04 | 2015-07-20 | Gas turbine engine blade tip treatment |
Publications (2)
Publication Number | Publication Date |
---|---|
US20160032738A1 true US20160032738A1 (en) | 2016-02-04 |
US10094227B2 US10094227B2 (en) | 2018-10-09 |
Family
ID=55179529
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US14/803,497 Active 2037-04-13 US10094227B2 (en) | 2014-08-04 | 2015-07-20 | Gas turbine engine blade tip treatment |
Country Status (1)
Country | Link |
---|---|
US (1) | US10094227B2 (en) |
Cited By (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160177732A1 (en) * | 2014-07-22 | 2016-06-23 | United Technologies Corporation | Hollow fan blade for a gas turbine engine |
EP3282090A1 (en) * | 2016-08-12 | 2018-02-14 | Hamilton Sundstrand Corporation | Airfoil systems and methods of assembly |
CN109321865A (en) * | 2018-12-06 | 2019-02-12 | 江苏丰东热技术有限公司 | One kind forming MoSi in titanium alloy surface2The method of antioxidant coating |
CN114107916A (en) * | 2022-01-26 | 2022-03-01 | 北京航空航天大学 | Plating method for keeping blade air film cooling hole smooth |
US20230128806A1 (en) * | 2021-10-27 | 2023-04-27 | General Electric Company | Airfoils for a fan section of a turbine engine |
Citations (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6159618A (en) * | 1997-06-10 | 2000-12-12 | Commissariat A L'energie Atomique | Multi-layer material with an anti-erosion, anti-abrasion, and anti-wear coating on a substrate made of aluminum, magnesium or their alloys |
US20070099027A1 (en) * | 2005-10-28 | 2007-05-03 | Anand Krishnamurthy | Wear resistant coatings |
US7285312B2 (en) * | 2004-01-16 | 2007-10-23 | Honeywell International, Inc. | Atomic layer deposition for turbine components |
US7942638B2 (en) * | 2005-06-29 | 2011-05-17 | Mtu Aero Engines Gmbh | Turbomachine blade with a blade tip armor cladding |
US20130004309A1 (en) * | 2011-03-02 | 2013-01-03 | Applied Thin Films, Inc. | Protective internal coatings for porous substrates |
US8470458B1 (en) * | 2006-05-30 | 2013-06-25 | United Technologies Corporation | Erosion barrier for thermal barrier coatings |
US20140010663A1 (en) * | 2012-06-28 | 2014-01-09 | Joseph Parkos, JR. | Gas turbine engine fan blade tip treatment |
US8668447B2 (en) * | 2008-12-26 | 2014-03-11 | Kabushiki Kaisha Toshiba | Steam turbine blade and method for manufacturing the same |
US20140272310A1 (en) * | 2013-03-15 | 2014-09-18 | Rolls-Royce Corporation | Coating interface |
US20150140331A1 (en) * | 2011-10-18 | 2015-05-21 | University Of Georgia Research Foundation, Inc. | Nanoparticles and method of making nanoparticles |
US20150290771A1 (en) * | 2012-03-27 | 2015-10-15 | Yundong Li | Abrasive article and method for making the same |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20130149163A1 (en) | 2011-12-13 | 2013-06-13 | United Technologies Corporation | Method for Reducing Stress on Blade Tips |
US9341066B2 (en) | 2012-06-18 | 2016-05-17 | United Technologies Corporation | Turbine compressor blade tip resistant to metal transfer |
-
2015
- 2015-07-20 US US14/803,497 patent/US10094227B2/en active Active
Patent Citations (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6159618A (en) * | 1997-06-10 | 2000-12-12 | Commissariat A L'energie Atomique | Multi-layer material with an anti-erosion, anti-abrasion, and anti-wear coating on a substrate made of aluminum, magnesium or their alloys |
US7285312B2 (en) * | 2004-01-16 | 2007-10-23 | Honeywell International, Inc. | Atomic layer deposition for turbine components |
US7942638B2 (en) * | 2005-06-29 | 2011-05-17 | Mtu Aero Engines Gmbh | Turbomachine blade with a blade tip armor cladding |
US20070099027A1 (en) * | 2005-10-28 | 2007-05-03 | Anand Krishnamurthy | Wear resistant coatings |
US8470458B1 (en) * | 2006-05-30 | 2013-06-25 | United Technologies Corporation | Erosion barrier for thermal barrier coatings |
US8668447B2 (en) * | 2008-12-26 | 2014-03-11 | Kabushiki Kaisha Toshiba | Steam turbine blade and method for manufacturing the same |
US20130004309A1 (en) * | 2011-03-02 | 2013-01-03 | Applied Thin Films, Inc. | Protective internal coatings for porous substrates |
US20150140331A1 (en) * | 2011-10-18 | 2015-05-21 | University Of Georgia Research Foundation, Inc. | Nanoparticles and method of making nanoparticles |
US20150290771A1 (en) * | 2012-03-27 | 2015-10-15 | Yundong Li | Abrasive article and method for making the same |
US20140010663A1 (en) * | 2012-06-28 | 2014-01-09 | Joseph Parkos, JR. | Gas turbine engine fan blade tip treatment |
US20140272310A1 (en) * | 2013-03-15 | 2014-09-18 | Rolls-Royce Corporation | Coating interface |
Cited By (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160177732A1 (en) * | 2014-07-22 | 2016-06-23 | United Technologies Corporation | Hollow fan blade for a gas turbine engine |
EP3282090A1 (en) * | 2016-08-12 | 2018-02-14 | Hamilton Sundstrand Corporation | Airfoil systems and methods of assembly |
US10815797B2 (en) | 2016-08-12 | 2020-10-27 | Hamilton Sundstrand Corporation | Airfoil systems and methods of assembly |
CN109321865A (en) * | 2018-12-06 | 2019-02-12 | 江苏丰东热技术有限公司 | One kind forming MoSi in titanium alloy surface2The method of antioxidant coating |
US20230128806A1 (en) * | 2021-10-27 | 2023-04-27 | General Electric Company | Airfoils for a fan section of a turbine engine |
CN114107916A (en) * | 2022-01-26 | 2022-03-01 | 北京航空航天大学 | Plating method for keeping blade air film cooling hole smooth |
Also Published As
Publication number | Publication date |
---|---|
US10094227B2 (en) | 2018-10-09 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US10094227B2 (en) | Gas turbine engine blade tip treatment | |
US9988913B2 (en) | Using inserts to balance heat transfer and stress in high temperature alloys | |
US10406596B2 (en) | Core arrangement for turbine engine component | |
US10030532B2 (en) | Abradable seal with thermally conductive microspheres | |
US20140010663A1 (en) | Gas turbine engine fan blade tip treatment | |
US10215030B2 (en) | Cooling hole for a gas turbine engine component | |
US10167738B2 (en) | Compressor case snap assembly | |
US20130340406A1 (en) | Fan stagger angle for geared gas turbine engine | |
US20230304410A1 (en) | System and Method for Manufacture of Abrasive Coating | |
US10041358B2 (en) | Gas turbine engine blade squealer pockets | |
US10822971B2 (en) | Cooling hole for a gas turbine engine component | |
EP2959115B1 (en) | Abradable seal including an abradability characteristic that varies by locality | |
US20160298463A1 (en) | Enhanced cooling for blade tip | |
US9376920B2 (en) | Gas turbine engine cooling hole with circular exit geometry | |
US10539036B2 (en) | Abradable seal having nanolayer material | |
US20160177734A1 (en) | Airfoil showerhead pattern apparatus and system | |
US20140272166A1 (en) | Coating system for improved leading edge erosion protection | |
US20160169000A1 (en) | Heat transfer pedestals with flow guide features | |
US20160084167A1 (en) | Self-modulated cooling on turbine components | |
BR102015032102A2 (en) | engine component and methods of producing an engine component and repairing an engine component | |
US20200182070A1 (en) | Internal cooling cavity with trip strips | |
US20170074116A1 (en) | Method of creating heat transfer features in high temperature alloys | |
US9957814B2 (en) | Gas turbine engine component with film cooling hole with accumulator | |
EP2938856A1 (en) | Gas turbine engine component cooling arrangement | |
US11149548B2 (en) | Method of reducing manufacturing variation related to blocked cooling holes |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:DING, ZHONGFEN;GUO, CHANGSHENG;JAWOROWSKI, MARK R.;REEL/FRAME:036134/0601 Effective date: 20140729 |
|
STCF | Information on status: patent grant |
Free format text: PATENTED CASE |
|
AS | Assignment |
Owner name: RAYTHEON TECHNOLOGIES CORPORATION, MASSACHUSETTS Free format text: CHANGE OF NAME;ASSIGNOR:UNITED TECHNOLOGIES CORPORATION;REEL/FRAME:054062/0001 Effective date: 20200403 |
|
AS | Assignment |
Owner name: RAYTHEON TECHNOLOGIES CORPORATION, CONNECTICUT Free format text: CORRECTIVE ASSIGNMENT TO CORRECT THE AND REMOVE PATENT APPLICATION NUMBER 11886281 AND ADD PATENT APPLICATION NUMBER 14846874. TO CORRECT THE RECEIVING PARTY ADDRESS PREVIOUSLY RECORDED AT REEL: 054062 FRAME: 0001. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF ADDRESS;ASSIGNOR:UNITED TECHNOLOGIES CORPORATION;REEL/FRAME:055659/0001 Effective date: 20200403 |
|
MAFP | Maintenance fee payment |
Free format text: PAYMENT OF MAINTENANCE FEE, 4TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1551); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY Year of fee payment: 4 |
|
AS | Assignment |
Owner name: RTX CORPORATION, CONNECTICUT Free format text: CHANGE OF NAME;ASSIGNOR:RAYTHEON TECHNOLOGIES CORPORATION;REEL/FRAME:064714/0001 Effective date: 20230714 |